Apparatus for testing the principles of detonation combustion



Dec. 28, 1965 s N LE JR" ET AL 3,225,589

APPARATUS FOR TESTING THE PRINCIPLES OF DETONATION COMBUSTION FiledApril 10, 1.961

INVENTORS SELDE/V B. SPANGLER JR. BY JAMES L. I??? ATTORNEY UnitedStates Patent Ofiice 3,225,589 Patented Dec. 28, 1965 APPARATUS FORTESTING THE PRINCIPLES OF DETONATIGN CQMBUSTION Selden B. Spangler, Jr.,and James L. Munier, Scottsdale,

Ariz., assignors to The Garrett Corporation, Los Angeles, Calif., acorporation of California Filed Apr. 10, 1961, Ser. No. 101,890

9 Claims. (Cl. 73-35) This invention relates to test apparatus, and moreparticularly to an apparatus for simulating the air flow and highaltitude conditions of hypersonic flight for the purpose of testing theprinciples of detonation combustion in hypersonic aircraft.

Recently, considerable work has been done on the development of avariable geometry detonation combustion engine for use in hypersonicaircraft. Such an engine and its theory and method of operation aredisclosed in the copending application of Hunter and Norman, Serial No.88,149, filed February 9, 1961. combustion process as described byHunter and Norman, the flame front or detonation is established andmaintained in a variable geometry aerothermodynamic duct or engine by ashock wave which has the characteristic that a static temperature riseoccurs in a very short distance across the wave. The detonation isstabilized in the engine by causing the shock wave to be maintained in afixed position relative to the confining structure, and subsequentexpansion of the gaseous detonation products is used to develop acontinuous thrust for the hypersonic aircraft.

In order to study the various aspects of the detonation combustionprocess and the many variables encountered in supersonic and hypersonicflight, it is desirable to have a test tunnel in which the anticipatedoperating conditions may be simulated. While work has been done on thedetonation combustion process in test tunnels, any such tunnels thathave been available in the past have had certain deficiencies. Forexample, prior detonation combustion tunnels have not been designed toreproduce flight conditions for which the detonation combustion processhas been shown to have most promise. In particular, the combustion airtotal temperature conditions have not been simulated; nor have fuelinjection conditions been simulated. In addition, oxygen from thecombustion air has been used in conjunction with a fuel for preheatingthe combustion air stream, with the result that there is a deficiency ofoxygen in the detonation combustion zone. The present invention isbased, at least in part, on the discovery that when hydrogen is used asthe fuel in a test tunnel for the detonation combustion process, mostprior difliculties may be obviated by preheating the combustion airstream with a flame derived from controlled outside sources of hydrogenand oxygen. In addition, preignition of the detonation fuel may beprevented and fuel injection more adequately studied by adding the fuelat a point or zone in the supersonic portion of the air stream at whichthe static temperature is below the ignition temperature of hydrogen inair. It has also been found that the overall size of the entireapparatus may be minimized by utilizing bleed air from a gas turbine asthe source of high pressure combustion air.

It is an object of this invention to provide an improved apparatus forsimulating the flow and altitude conditions of hypersonic flight fortesting detonation combustion.

Another object of this invention is to provide a test apparatus forsimulating hypersonic flight fluid flow for detonation combustion inwhich a stream of combustion air is supplied from a high speed gasturbine.

Another object of this invention is to provide a test apparatus forsimulating hypersonic flight fluid flow for detonation combustion inwhich the combustion air In the detonation d stream is preheated with agas flame derived from controlled outside gas sources so as to maintaina predetermined oxygen supply in the air stream.

A further object of this invention is to provide a test apparatus forsimulating hypersonic flight fluid flow for detonation combustion inwhich fuel is added in the supersonic portion of the air stream wherethe static temperature is below the ignition temperature of hydrogen inair.

It is another object of this invention to provide a test apparatus forsimulating hypersonic flight fluid flow for detonation combustion inwhich the fuel injector is streamlined so as to reduce any disturbancein the air stream to a minimum.

A still further object of this invention is to provide a test apparatusfor simulating hypersonic flight fluid flow for detonation combustion inwhich the geometry and location of the preheater and fuel injector inthe apparatus are such as to provide the proper temperature and fueldistribution over an appreciable portion of the detonation combustionzone, which portion does not extend to the walls of the apparatus sothat there will be no deleterious effects on the walls due to hightemperature.

The above and other features and objects of the invention will beapparent from the following description and the accompanying drawing, inwhich:

FIG. 1 is a schematic side elevational view of a supersonic testingapparatus embodying the features of this invention;

FIG. 2 is a longitudinal vertical sectional view of portions of thepreheater and tunnel sections of the apparatus;

FIG. 3 is a horizontal sectional view of the preheater burner, takenalong the line 33 of FIG. 2; and

FIG. 4 is a horizontal sectional view of the fuel injector, taken alongthe line 4-4 of FIG. 2.

Referring now to the drawing, and particularly FIG. 1, it will beobserved that the test apparatus in which the features of this inventionhave been incorporated comprises a source of high pressure air 10 fromwhich the air is directed through a stilling chamber 11 where itsvelocity is reduced, turbulence eliminated, and the air stream isrendered substantially homogeneous as it leaves said chamber. Thestilled air stream then passes through a preheater 12 to an inlet 13 ofa test tunnel or elongated aerothermodynamic duct 14 having an exhaust15. In the present instance, the duct 14 is shown as rectangular incross section, though obviously it could have any desired cross section,such as polygonal, elliptical or circular. The test apparatus isdesigned particularly for simulating the operating conditions of adetonation combustion engine in hypersonic flight. While any convenientsource of high pressure air may be used, it has been found to beadvantageous to utilize the high pressure bleed air stream produced by agas turbine 16, since this reduces the size and weight of the overallapparatus. As is well known, such turbines usually include a compressorsection 17 adjoining a combustor section 18 which leads to the turbine16, the burned gases being discharged through an exhaust 29. In thisinstance, a conduit 21 connected to the compressor section 17 leads thehigh pressure bleed air stream to the stilling chamber 11 wherein itsvelocity is reduced and flow inhomogeneities are minimized. Thesubstantially homogeneous flow is then ducted to the preheater 12. Inthe event that electric power may be needed in the test apparatus, suchas for operating thermocouples and other test instruments, the gasturbine 16 could obviously be used to drive a suitable electricgenerator (not shown).

Prior to the present invention, there has been some question, in the useof test tunnels of this general character, as to the variation of oxygenconcentration in the detonation combustion zone and the concomitanteffect of such variation on the characteristics of the combustionprocess because of the use of the oxygen in the combustion air incombination with a fuel to preheat the air. According to this invention,the air stream is preheated with a flame produced by independent andseparately controlled sources of hydrogen and oxygen. As shown in FIGS.2 and 3, the preheater 12 includes a housing 22 to which the air conduit21 is connected, and air from the compressor 17 thus passes through thepreheater 12 into the inlet 13 of the test tunnel. A preheater burner 23is suitably mounted in the housing and provided with a starter section24 which includes a glow-plug 25 connected to a convenient source ofpower 26. The burner 23, in the form shown, comprises a streamlined orwedgeshaped upstream end 27 having a relatively sharp leading edge 28 onthe upstream side thereof and a vertically disposed groove 30 on thedownstream side or edge thereof. A central burner opening 31 ispositioned in the groove 3-9 in a vertical position correspondingapproximately to the center of the air stream, and this opening isconnected by a passage 32 and line or conduit 33 to a pressure source ortank 34 of hydrogen gas. A valve 35 downstream of a pressure regulator36 in the line 33 may be used to control the flow of fuel to the burneropening 31. Oxygen required for combustion of the fuel is directed tothe burner area through a plurality of openings 37 which may be arrangedin a circle around the opening 31. The openings 37 are connected throughpassages 38 to a conduit or pipe line 39 which leads to a source or tank40 of oxygen under the control of a flow control valve 41 and a pressureregulator 42. It will be understood that, if desired, the gases could bereversed in the burner so that oxygen would be supplied to the opening31 and hydrogen to the surrounding openings 37.

The starter section 24 has connected to it a conduit 43 which leads to amixer 44. Hydrogen and oxygen are supplied to the mixer 44 by conduits45 and 46, respectively, which emanate from the tanks 34 and 40.Pressure regulator 47 and flow control valve 48 regulate the hydrogenflow to the mixer 44; and similarly pressure regulator 5t) and flowcontrol valve 51 regulate the flow of oxygen from tank 40 to said mixer.When it is desired to start the preheater burner 23, hydrogen and oxygenare first supplied to the mixer 44 and the mixed gases flow throughconduit 43 to the igniter section 24. The glowplug 25 is supplied withelectric current and causes the gas mixture flowing into groove 30 fromconduit 43 to be ignited. The flame moves along the groove to theopenings 31 and 37 which are then supplied with hydrogen and oxygenthrough conduits 33 and 39 and these gases mix and are ignited. Theglow-plug 25 and igniter gas flow may then be turned off.

By having a separate supply of oxygen for the preheater, as described,proper amounts of oxygen may be supplied to the burner 23 so that theproducts of combustion, which are water vapor and excess oxygen, willhave the same oxygen-to-inerts (water vapor) ratio as that or"oxygen-to-inerts (nitrogen plus traces of other noncombustibles) innormal unburned air. In this manner, the preheat process will have theleast effect on the oxygen concentration at the detonation combustionzone. In addition, the oxygen is most thoroughly mixed in the air-watervapor stream because of its direct use in the preheat combustionprocess. The design and operation of the preheater are such that a largeproportion of the air stream entering the test tunnel is heated topredetermined temperatures as high as 3500 F.

Detonation combustion, which will produce optimum thrust conditions forhypersonic flight in the neighborhood of Mach 6.5, requires supersonicair flow in the detonation combustion zone of about Mach 3 and atemperature of 3400 F. To increase the velocity of the heated air streamas it flows from the preheater into the tunnel inlet 13, a nozzle 52having a restricted throat 53 is provided in the tunnel 14 adjacent saidinlet 13. The internal geometry of the tunnel also includes specialwedges 54 on the top and bottom inside surfaces toward the downstreamend of the tunnel and these act to hold a normal shock wave 55 in thedesired position in the tunnel, as explained in the Hunter and Normanapplication referred to above. It is across this normal shock wave thatdetonation combustion takes place when fuel (in this instance, hydrogen)is supplied to the tunnel.

A special streamlined fuel injector 56 is provided to supply hydrogenfuel to the tunnel and inject such fuel into the supersonic air stream.This injector, which in the form shown is diamond-shaped, comprises anupstream Wedge-shaped portion 57 having a sharp leading edge 58 and anadjoining downstream wedge-shaped section 60. These two wedge-shapedsections having their bases contiguous to one another provide thedesired streamlined shape for the injector 56, which is shown asvertically disposed substantially midway between the sides of the ductor tunnel in a position downstream of the nozzle 52 in a uniform flowfield but upstream of the shock wave 45. It will be understood that theinjector could be horizontally disposed midway between the top andbottom of the tunnel, if desired. Fuel may be injected into the passingsupersonic air stream through an opening or nozzle 61 provided in thedownstream edge 62 of the injector. Nozzle 61 is connected by a passage63 and a conduit 64 to the hydrogen tank 34 and is under the control ofa flow control valve 65 and a pressure regulator 66.

To cool the fuel injector, and thus aid in preventing preignition of thefuel, a cooling fluid inlet passage 67 is formed in the upstreamwedge-shaped portion 57 adjacent the leading edge 58. This passage isconnected by a conduit 68 to a source 70 (FIG. 1) of coolant. Passage 67is connected with a centrally disposed return passage 71, which in turnis connected to a conduit 72 leading to a sump (not shown) to collectthe used cooling fluid. A pump 73 and valve 74 control the flow ofcooling fluid, which in this instance is water, from the tank or source70 to the fuel injector cooling passages 67 and 71 and also to a conduit75. This latter conduit leads to a cooling header 76 disposed over thelength of the tunnel 13 and provided with a plurality of sprinklernozzles 77. In this way the entire tunnel may be water-cooled.

In the event that it is desired to cool the tunnel nozzle 52, coolingpassages 78 may be provided around the throat 53, as shown in FIG. 2. Apipe or conduit 80 is then connected to such cooling passages and to thesource of cooling fluid. In this instance, the conduit 80 is shown inFIG. 1 as being connected to the header 76.

Since one purpose of the apparatus is to study the detonation wave,observation windows 81 may be provided on each side of the tunnel 14 ata position adjacent the wedges 54.

By means of the test apparatus described above, it is possible to studythe many variables that are involved in the detonation combustionprocess, and also to study such variables and conditions as they may beencountered in a variable geometry engine under changing flightconditions. One of the principal and important problems in such anengine is the method of injecting and controlling the fuel flow so as toeffect minimum specific fuel consumption. Consequently, the regulationand control of the oxygen in the air stream to simulate air conditionsat flight altitude is important and according to this invention ispossible with the specially controlled preheater.

Various changes may be made in the construction and certain features maybe employed without others Without departing from the invention orsacrificing any of its advantages.

We claim:

.1. A test apparatus for simulating the air flow conditions inhypersonic flight for testing the principles of detonation combustion,comprising: an elongated aerothermodynamic duct; means for supplying airat high pressure to said duct; means for preheating such air while atthe same time regulating the amount of oxygen therein; means forestablishing and maintaining a shockwave in said duct; and means forinjecting fuel into the airstream in said duct so that detonation maytake place across such shockwave.

2. A test apparatus for simulating the air flow conditions in hypersonicflight for testing the principles of detonation combustion, comprising:an elongated aerothermodynamic duct; a gas turbine for supplying an airstream at high pressure to said duct; means between said turbine andduct for preheating said air stream while at the same time controllingthe amount of oxygen in said air stream; means for establishing andmaintaining a shockwave in said duct; and means for injecting fuel intothe airstream in said duct so that detonation may take place across suchshockwave.

3. A test apparatus for simulating the air flow conditions in hypersonicflight for testing the principles of detonation combustion, comprising:an elongated aerothermodynamic duct; means for supplying air at highpressure to said duct; means for preheating such air while at the sametime regulating the amount of oxygen therein; means for creating anormal shock wave in said duct; and means upstream from said shock Wavefor injecting controlled amounts of fuel into the preheated supersonicair stream.

4. A test apparatus for simulating the air flow conditions in hypersonicflight for testing the principles of detonation combustion, comprising:an elongated aerothermodynamic duct; means for supplying an air streamat high pressure to said duct; means for preheating said air streamwhile at the same time regulating the amount of oxygen therein; meansfor accelerating said air stream to supersonic velocities; andstreamlined means in said duct for injecting controlled amounts of fuelinto the preheated air stream after acceleration of said air stream tosupersonic velocities.

5. A test apparatus for simulating the air flow conditions in hypersonicflight for testing the principles of detonation combustion, comprising:an elongated aerothermodynamic duct; means for supplying an air streamat high pressure to said duct; means for preheating said air streamWhile at the same time regulating the amount of oxygen therein; meansfor accelerating said air stream to supersonic velocities; streamlinedmeans in said duct for injecting controlled amounts of fuel into thepreheated air stream after acceleration thereof to supersonicvelocities; and means for circulating cooling fluid through such fuelinjecting means.

6. A test apparatus for simulating supersonic air flow and high altitudeconditions for testing the principles of detonation combustion,comprising: an elongated aerothermodynamic duct; means for supplying anair stream at high pressure to said duct; means for stilling said airstream; means for preheating said air stream while at the same timeregulating the amount of oxygen therein; means for accelerating said airstream to supersonic velocities; streamlined means in said duct forinjecting controlled amounts of fuel into the preheated air stream afteracceleration thereof to supersonic velocities; and means for cooling theoutside of said duct and the inside of such fuel injecting means.

7. A test apparatus for simulating supersonic air flow and high altitudeconditions for testing the principles of detonation combustion,comprising: an elongated aerothermodynamic duct; means for supplying anair stream at high pressure to said duct; means for stilling said airstream; means for preheating said air stream While at the same timeregulating the amount of oxygen therein; means for accelerating said airstream to supersonic velocities; streamlined means in said duct forinjecting controlled amounts of fuel into the preheated air stream afteracceleration thereof to supersonic velocities; and means for cooling theair stream accelerating means and the inside of such fuel injectionmeans.

8. A test apparatus for simulating the air flow conditions in hypersonicflight for testing the principles of detonation combustion, comprising:an elongated aerothermodynamic duct; means for supplying air at highpressure to said duct; means for preheating such air while at the sametime regulating the amount of oxygen therein; means for creating anormal shock Wave in said duct; and means upstream from said shock Wavefor injecting controlled amounts of fuel into the preheated supersonicair stream, the geometry and location of the preheating means and thefuel injecting means being such as to provide proper temperature andfuel distribution over an appreciable portion of said normal shock waveWhere detonation combustion takes place, such portion being of limitedextent so as to prevent harm to the walls of the duct due to hightemperature.

9. A test apparatus for simulating the air flow conditions in hypersonicflight for testing the principles of detonation combustion, comprising:an elongated aerothermodynamic duct; means for supplying air at highpressure to said duct; means for preheating such air, said preheatingmeans including means for supplying fuel thereto and supplying regulatedamounts of oxygen to said preheating means and said duct; means forcreating a normal shockwave in said duct; and means upstream from saidshockwave for injecting controlled amounts of fuel into the preheatedsupersonic air stream.

References Cited by the Examiner UNITED STATES PATENTS 2,615,331 10/1952Lundgren 73-116 2,763,155 9/1956 Beams et al. 73116 2,885,890 5/1959Liccini et al. 73--147 3,005,338 10/1961 Libby et al 73147 OTHERREFERENCES Cushman, R. H; Aviation Week, January 6, 1958, pages 57, 58,59, 63.

Dugger, G. L.: ARS Journal, November 1959, pages 819-827.

LOUIS R. PRINCE, Primary Examiner.

DAVID SCHONBERG, Examiner.

1. A TEST APPARATUS FOR SIMULATING THE AIR FLOW CONDITIONS IN HYPERSONICFLIGHT FOR TESTING THE PRINCIPLES OF DETEONATION COMBUSTION, COMPRISING:AN ELONGATED AEROTHERMODYNAMIC DUCT; MEANS FOR SUPPLYING AIR AT HIGHPRESSURE TO SAID DUCT; MEANS FOR PREHEATING SUCH AIR WHILE AT THE SMETIME REGULATING THE AMOUNT OF OXYGEN THEREIN; MEANS FOR ESTABLISHING ANDMAINTAINING A SHOCKWAVE IN SAID DUCT; AND MEANS FOR INJECTING FUEL INTOTHE AIRSTREAM IN SAID DUCT SO THAT DETONATION MAY TAKE PLACE ACROSS SUCHSHOCKWAVE.